This invention relates generally to cooling systems and, more particularly, to a cooling system for use in high temperature gas turbines.
Cooling of high temperature components in gas turbine engines is one of the most challenging problems facing engine designers today, particularly as it relates to the turbine portions of the engine where temperatures are most severe. While improved high temperature materials have been developed which partially alleviate the problem, it is clear that complete reliance on advanced technology materials will not be practical for the foreseeable future. One reason is that these advanced materials contemplate expensive manufacturing techniques or comprise alloys of expensive metals. Thus, the product, though technically feasible, may not be cost effective. Additionally, as gas turbine temperatures are increased to higher and higher levels, it is clear that no contemplated material, however exotic, can withstand such an environment without the added benefit of supplemental fluid cooling. Fluid cooling, therefore, can permit the incorporation of more cost effective materials into present-day gas turbine engines and will permit the attainment of much higher temperatures (and, therefore, more efficient engines) in the future.
One area of the turbine which is particularly troublesome in this regard is the turbine nozzle band which comprises a number of annular sectors which form a complete circular wall to define a flow path for the operating fluid of the turbomachine through a stage of turbine nozzle vanes. Various fluid cooling techniques have been proposed in the past to cool these band sectors, these various techniques being commonly classified as convection, impingement and film cooling. All of these methods have been tried, both individually and in combination, utilizing the relatively cool pressurized air from the compressor portion of the engine as the coolng fluid. Such prior art concepts are discussed in U.S. Pat. No. 3,800,864, issued to Ambose A. Hauser et al, which is assigned to the assignee of the present invention. Although these various prior art approaches toward cooling turbine nozzle bands are structurally distinguishable, these designs all remove heat in substantially the same manner. That is to say, they all appear to incorporate backside heat convection cavities. These cavities are generally formed by brazing a back plate to the nozzle band, with cooling accomplished either by impinging a fluid coolant from a coolant plenum through the plate and onto the backside of the band or by passing a coolant over a multiplicity of pin fins extending between the band and the plate, thereby heating the coolant and cooling the band. This spent coolant is then dumped as a film over the band hot surface.
Such systems, while basically effective in providing turbine band cooling in many turbine configurations, have several shortcomings. Most importantly, since the turbine band comprises a number of distinct annular sectors which abut each other to form a circular wall, leakage of the pressurized cooling air occurs through the gaps between the ends of adjacent sectors since the driving pressure for impingement cooling of the bands is the same driving pressure for gap leakage. This is particularly true when the cooling air supply pressure is high compared to motive gas stream pressures, such as in the low pressure turbine section of a gas turbofan engine.
Another unsatisfactory characteristic of prior art systems is that they are very expensive to manufacture and difficult (and costly) to repair. Typically, they contemplate intricate castings characterized by a multiplicity of pin fins or coolant passages, or comprise a cast shroud sector to which is brazed a perforated impingement liner which combine to form a single plenum for cooling. While the impingement cooling systems may require less complex castings than their convection-cooled counterparts, they suffer from the disadvantage that particulate matter may become lodged within the liner perforations and substantially reduce the cooling effectiveness. This requires replacement of the entire band sector.
Furthermore, it is a costly and time-consuming process to tune the impingement cooling systems in new turbine designs. Since the band may be subjected to localized heat concentration such as hot streaks, either sufficient air must be supplied to the entire plenum between the band sector and the impingement liner to cool in the area of these hot streaks (i.e., the coolant flow is established by the portion of the sector subjected to the highest temperature --clearly a waste of coolant) or else different impingement liner perforation patterns must be tried, each of which must be brazed to the sector castings. This is a time-consuming process.
Finally, these characteristics are compounded many fold when the turbine is of the variable area variety, employing turbine vanes which are rotatable about their longitudinal axes and which protrude through the annular band sectors. In these turbines, the vane trunnions penetrate the coolant supply cavities, presenting difficult problems of coolant routing and sealing. For example, the gap between variable vane trunnions and band sectors offers another source of coolant leakage.
Minimization of coolant leakage is important since the source of turbine coolant is usually air that is bled from the compressor portion of the engine and, as such, has had work done on it by the compressor. However, since leakage air loses much of its pressure as it flows through the gap and cracks, it does not return its full measure of work to the propulsive cycle. Additionally, the reintroduction of cooling air into the gas stream produces a loss in gas stream total pressure as a result of momentum mixing losses associated with injecting relatively low total pressure cooling air into a high total pressure gas stream. Thus, the greater the amount of cooling air which is lost through leakage, the greater the propulsive cycle efficiency losses become. It will, therefore, be appreciated that a cooling system which reduces leakage between adjacent band sectors will result in a more efficient turbine.